1. Field of the Invention
This invention relates to film cooled combustor liners for use in gas turbine engines, and more particularly, to aircraft gas turbine engine multi-hole film cooled combustor liners.
2. Description of Related Art
Combustor liners are generally used in the combustion section of a gas turbine engine which is located between the compressor and turbine sections of the engine. Combustor liners are also used in the exhaust sections of aircraft engines that employ afterburners. Combustors generally include an exterior casing and an interior combustor. Fuel is burned in the interior of the combustor producing a hot gas usually at an intensely high temperature such as 3,000.degree. F. or even higher. To prevent this intense heat from damaging the combustor before it exits to a turbine, a heat shield or combustor liner is provided in the interior of the combustor. This combustor liner thus prevents the intense combustion heat from damaging the combustor or surrounding engine.
Some aircraft gas turbine engines, particularly ones that are capable of supersonic flight such as military fighters and bombers, have afterburners or augmenters located in the exhaust section of the engine. Heat shields or liners are also provided for afterburner engines to prevent the intense combustion heat from damaging the surrounding casing of the exhaust section or other parts of the engine and aircraft.
In the past, various types of combustor and afterburner liners have been suggested and used. In addition, a variety of different methods have been suggested how to cool these liners so as to withstand greater combustion heat or prolong the usable life expectancy of the liner. Characteristically these liners are excessively complex, difficult and expensive to manufacture and overhaul, and increase the weight of the engine. Engine designers have long sought to incorporate low weight liners capable of withstanding the temperatures and pressure differentials found in combustors that are relatively easy and inexpensive to manufacture.
Prior methods for film cooling combustion liners provided circumferentially disposed rows of film cooling slots such as those depicted in U.S. Pat. No. 4,566,280 by Burr and U.S. Pat. No. 4,733,538 by Vdoviak et al. which are typified by complex structures that have non-uniform liner thicknesses which give rise to thermal gradients which cause low cycle fatigue in the liner and therefore shorten their potential life expectancy and reduce their durability. The complex shapes and machining required to produce these liners negatively effects their cost and weight.
Other film cooled combustor liners, such as those depicted in U.S. Pat. No. 4,695,247 by Enzaki et al., have disclosed the use of double wall liners which employ film cooling holes having about 30 degrees incline to the film cooled hot wall. This type of double walled liner is complex, heavy and expensive to manufacture and repair. The interior cavity between the spaced apart double walls can also cause maintenance problems and clogging. Reference is also made to another double wall type liner having multi hole film cooling disclosed in U.S. Pat. No. 4,896,510 by Foltz. Foltz is cited for reference purposes only in order to better understand the present invention and therefore it, as well as all the other patents above, are specifically incorporated herein by reference.
Yet another means to cool liners employs transpiration cooling means wherein cooling air is continuously effused through a liner which can almost be described as porous. Transpiration cooled liners have relatively complicated holes and may employ multiple layers of material. Transpiration cooled liners may be flimsy because of the degree of porosity and, therefore, require complicated or heavy support means in order to be used in modern day large turbine engines.